Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity

ABSTRACT

An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a ratio of a number of stages in the low-pressure turbine to a number of stages in the high-pressure turbine is 2.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is continuation of U.S. patent application Ser. No.13/163,304, which was filed on 17 Jun. 2011. U.S. patent applicationSer. No. 13/163,304 is a divisional of U.S. Pat. No. 7,984,607, whichwas filed 6 Sep. 2007. U.S. patent application Ser. No. 13/163,304 andU.S. Pat. No. 7,984,607 are both hereby incorporated by reference.

BACKGROUND

1. Technical Field

The disclosure generally relates to gas turbine engines.

2. Description of the Related Art

Turbine durability for high inlet temperature/high fuel-to-air ratio gasturbine engine designs can be compromised by gas path temperature and/orchemical species non-uniformities occurring at the exit of a combustor.Chemical species variation is particularly relevant to high fuel-to-airratio turbine designs in that the combustion process may not becompleted at the exit plane of the combustor section with the resultthat partial products of reaction (PPR's) enter the turbine section. Theanticipated presence of such temperature non-uniformities contributes tothe use of conservative cooling designs for the turbine in order toprevent damage that can be caused by a failure to account for suchtemperature/PPR's non-uniformities.

SUMMARY

A gas turbine engine according to an exemplary embodiment of the presentdisclosure includes, among other possible things, a turbine sectionoperative to impart rotational energy to a compressor section. Theturbine section includes at least a low-pressure turbine and ahigh-pressure turbine. A ratio of a number of stages in the low-pressureturbine to a number of stages in the high-pressure turbine is 2.

In a further non-limiting embodiment of the foregoing gas turbine engineembodiment, the low-pressure turbine has exactly four stages.

In a further non-limiting embodiment of either of the foregoing gasturbine engine embodiments, the high-pressure turbine has exactly twostages.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the turbine section has a first set of vanespositioned adjacent a combustion section and a first set of rotatableblades positioned downstream of and adjacent to the first set of vanes.A number of vanes of the first set of vanes exceeds a number of bladesof the first set of blades.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, a first stage of the high-pressure turbine comprisesthe first set of vanes and the first set of blades.

A gas turbine engine according to another exemplary embodiment of thepresent disclosure includes, among other possible things, a turbinesection operative to impart rotational energy to a compressor section.The turbine section includes at least a high-pressure turbine havingexactly two stages and a low-pressure turbine having exactly fourstages.

In a further non-limiting embodiment of the foregoing gas turbine engineembodiment, the high-pressure turbine has a first set of vanespositioned adjacent a combustion section and a first set of rotatableblades positioned downstream of and adjacent to the first set of vanes.A number of vanes of the first set of vanes exceeds a number of bladesof the first set of blades.

A method of operating a gas turbine engine according to an exemplaryaspect of the present disclosure includes, among other possible things,providing a gas turbine engine having a combustion section and a turbinesection. The turbine section has a first set of vanes and a first set ofblades. The first set of blades being located downstream from andadjacent to the first set of vanes. The turbine section includes ahigh-pressure turbine and a low-pressure turbine. The low-pressureturbine having four stages.

In a further non-limiting embodiment of the foregoing method ofoperating a gas turbine engine, the method includes completing acombustion reaction along a gas flow path prior to a plane defined bydownstream portions of the first set of blades.

In a further non-limiting embodiment of either of the foregoing methodsof operating a gas turbine engine, a vane-blade count ratio of theturbine stage is greater than unity, and completing the combustionreaction comprises mixing combustion products from the combustionsection using the vanes of the first set of vanes.

In a further non-limiting embodiment of any of the foregoing methods ofoperating a gas turbine engine, the engine is operative such thatcooling air provided to the first set of vanes reacts with combustionproducts from the combustion section in order to complete the combustionreaction.

In a further non-limiting embodiment of any of the foregoing methods ofoperating a gas turbine engine, the method includes film-cooling thevanes of the first set of vanes.

In a further non-limiting embodiment of any of the foregoing methods ofoperating a gas turbine engine, the gas turbine engine further comprisesa second set of vanes located downstream from the first set of blades,and the second set of vanes is film-cooled to a lesser extent than thatprovided to the first set of vanes.

DESCRIPTION OF THE FIGURES

Many aspects of the disclosure can be better understood with referenceto the following drawings. The components in the drawings are notnecessarily to scale. Moreover, in the drawings, like reference numeralsdesignate corresponding parts throughout the several views.

FIG. 1 is a schematic diagram depicting an embodiment of a gas turbineengine.

FIG. 2 is a schematic diagram depicting the embodiment of FIG. 1,showing detail of the first and second stages of the turbine section.

FIG. 3 is a flowchart depicting an embodiment of a method for operatinga gas turbine engine.

FIG. 4 is a flowchart depicting an embodiment of a method for designinga gas turbine engine.

DETAILED DESCRIPTION

Gas turbine engine systems and related methods involving vane-bladecount ratios greater than unity are provided, several exemplaryembodiments of which will be described. In this regard, an increasedvane count could increase the probability that a temperature/PPR'snonuniformity (“hot and/or fuel-rich streak”) will mix with cooler gasesand, therefore, dissipate before propagating beyond the rotating bladesof the first stage of the turbine. In some embodiments, the cooler gasesused for dissipating such a hot streak are provided as cooling air,which is provided for film-cooling the vanes of the first turbine stage.In some embodiments, the vane and blade counts deviate from a nominalnumber so that overall parasitic drag and weight directly attributableto the vanes and blades are comparable to a gas turbine enginecontaining an equal number of vanes and blades. For example, in a firststage turbine design incorporating forty-eight (48) vanes and sixty-two(62) blades, the number of vanes could be increased to fifty six(48+8=56), whereas the number of blades could be decreased to fifty four(62−8=54). It should be noted that increasing the number of vanes canpotentially enhance mixing of the gases departing the combustionsection, thereby reducing the requirement for conservative cooling ofdownstream vanes. Thus, an overall weight reduction may be achieved byreducing the requirements of cooling air in some embodiments.Additionally, aerodynamic efficiency may be improved as an increasednumber of vanes can yield higher levels of unsteady flow in thedownstream rotor passages. As such the formation of rotor passagesecondary flow vortices and losses can be inhibited.

Referring now in detail to the drawings, FIG. 1 is a schematic diagramdepicting an embodiment of a gas turbine engine 100. As shown in FIG. 1,engine 100 includes a compressor section 102, a combustion section 104and a turbine section 106. Notably, engine 100 is a turbofan although itshould be noted that the concepts described herein should not beconsidered limited to use with gas turbine engines configured asturbofans.

Turbine section 106 incorporates multiple stages, each of which includesa set of stationary vanes and a corresponding set of rotating blades. Inthis regard, a first stage 108 of the turbine section includes a firstset of vanes 110 and a first set of blades 112. The first stage of theturbine section is located immediately downstream of the combustionsection and immediately upstream of a second stage 114 of the turbine,which includes a second set of stationary vanes 116.

As shown in FIG. 2, blades 112 are located downstream of vanes 110,whereas vanes 116 are located downstream of blades 112. Notably,downstream portions of the blades 112 define an exit plane 120. Notably,interaction of gas 122 flowing along the gas path defined by the vanesand blades causes combustion products to mix and complete a combustionreaction prior to traversing the exit plane of the first set of blades112. This is accomplished, at least in part, by providing a greaternumber of vanes 110 than there are blades 112, i.e., the vane-bladecount ratio of the first turbine stage is greater than unity (1).

In the embodiment of FIGS. 1 and 2, vanes 110 and 116 incorporatefilm-cooling holes that direct cooling air for film-cooling the vanes.By way of example, vane 110 includes cooling holes 130, and vane 116includes cooling holes 132. Note that although the number of coolingholes in vane 110 exceeds the number of cooling holes in vane 116,various other numbers and arrangements of cooling holes can be providedin other embodiments.

Because of work extraction in a first turbine stage, the temperature ofgas is reduced at exit plane of that stage relative to the temperatureat the entrance of the first vanes. As a result, for conventionalturbine designs, cooling requirements for the downstream vanes areusually much lower than the cooling requirements for the first vanes.However, for high-fuel-to-air ratio turbine designs (i.e., designs thatexhibit significant PPR concentrations at the inlet to the first stage),the oxygen included in the cooling air provided in the first stage 108completes the combustion reaction and can significantly increase thetemperature of gas 122 temperature at exit plane 120. The degree ofcircumferential and radial uniformity of this hot and/or fuel-richstreak temperature increase is a factor that should be considered forthe cooling design of vanes 116. If the hot and/or fuel-rich streaks arehighly concentrated (non-uniform temperature at station 120), then allof the vanes 116 should be designed to accommodate the hottest possiblestreak.

However, if the hot and/or fuel-rich streaks are diffused by the firststage 108 (a highly uniform temperature at station 120), then the vanes116 can be designed to accommodate a lower peak temperature. This canresult in a weight reduction of the gas turbine engine as lighter and/orfewer components associated with routing of the cooling air may beprovided.

In this regard, FIG. 3 is a flowchart depicting an embodiment of amethod for operating a gas turbine engine. As shown in FIG. 3, themethod may be construed as beginning at block 302, in which a gasturbine engine having a combustion section and a turbine stage isprovided.

Notably, the turbine stage includes a first set of vanes and a first setof blades, with the first set of blades being located downstream fromand adjacent to the first set of vanes. In block 304, a combustionreaction is completed along a gas flow path prior to a plane defined bydownstream portions of the first set of blades. In some embodiments,design of such a first stage incorporates, through the use of both CFDanalysis and empirical correlations, any combination of (a)Vane(N)/Blade(N) ratio, (b) vane and/or blade film-cooling schemes or(c) vane and/or blade aerodynamic designs such that temperature andPPR's nonuniformities at the exit plane of the turbine stage arereduced, e.g., minimized or eliminated.

FIG. 4 is a flowchart depicting another embodiment of a method.Specifically, the flowchart of FIG. 4 involves a method for designing agas turbine engine. That method may be construed as beginning at block402, in which a number (N) of vanes and a corresponding number (N) ofblades for a turbine stage of a gas turbine engine are selected. Inblock 404, the number of the vanes is increased by M. In block 406, thenumber of blades is decreased by at least M such that the number ofvanes of the turbine stage exceeds the number of blades of the turbinestage. In some embodiments, the turbine stage is a first turbine stagedownstream of a combustor.

A set of numerical experiments were conducted to quantify the impact ofchange in the number of first vanes on the temperature non-uniformity atthe exit of a representative turbine stage. These simulations wereconducted by changing the number of airfoils in the vane row whileholding the airfoil count for the rotor row constant. The ratio of vanesto blades in these studies were 2/3, 111 and 3/2. The temperature atinlet to the stage was held constant at a typical combustor exittemperature value while the metal temperature for the vane wasmaintained at a constant value consistent with the airfoil durabilityrequirements. The rotor airfoils and endwalls, however, were maintainedat adiabatic wall temperatures. An unsteady 3-DReynoldsAveraged-Navier-Stokes CFD code was used to conduct thesesimulations. Results from these simulations indicated that the absolutetemperature distortion at the exit of the rotor was about 60%, 30% and16% of the inlet distortion to the rotor, which was constant for theabove three numerical experiments. These simulations clearly indicatethat increasing the number of vanes relative to the blades enhancesmixing between the hot and cold stream in the rotor passages.

Interrogation of numerical data from these simulations also indicatedthat the loss levels in the rotor passages were also reduced as thenumber of vanes was increased. In addition, increasing the number ofvanes was also found to reduce the hot spot temperature on the rotorairfoil pressure side indicating that the cooling air in the rotorpassages may also be favorably impacted by increased vane count.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. Thus, the scope of legal protectiongiven to this disclosure can only be determined by studying thefollowing claims.

1. A gas turbine engine comprising: a turbine section operative toimpart rotational energy to a compressor section, wherein the turbinesection includes at least a low-pressure turbine and a high-pressureturbine, and a ratio of a number of stages in the low-pressure turbineto a number of stages in the high-pressure turbine is
 2. 2. The gasturbine engine of claim 1, wherein the low-pressure turbine has exactlyfour stages.
 3. The gas turbine engine of claim 1, wherein thehigh-pressure turbine has exactly two stages.
 4. The gas turbine engineof claim 1, wherein the turbine section has a first set of vanespositioned adjacent a combustion section and a first set of rotatableblades positioned downstream of and adjacent to the first set of vanes,and a number of vanes of the first set of vanes exceeds a number ofblades of the first set of blades.
 5. The gas turbine engine of claim 4,wherein a first stage of the high-pressure turbine comprises the firstset of vanes and the first set of blades.
 6. A gas turbine enginecomprising: a turbine section operative to impart rotational energy to acompressor section, wherein turbine section includes at least ahigh-pressure turbine having exactly two stages and a low-pressureturbine having exactly four stages.
 7. The gas turbine engine of claim 6wherein the high-pressure turbine has a first set of vanes positionedadjacent a combustion section and a first set of rotatable bladespositioned downstream of and adjacent to the first set of vanes, whereina number of vanes of the first set of vanes exceeds a number of bladesof the first set of blades.
 8. A method for operating a gas turbineengine comprising: providing a gas turbine engine having a combustionsection and a turbine section, the turbine section having a first set ofvanes and a first set of blades, the first set of blades being locateddownstream from and adjacent to the first set of vanes, wherein theturbine section includes a high-pressure turbine and a low-pressureturbine, the low-pressure turbine having four stages.
 9. The method ofclaim 8, including completing a combustion reaction along a gas flowpath prior to a plane defined by downstream portions of the first set ofblades.
 10. The method of claim 8, wherein: a vane-blade count ratio ofthe turbine stage is greater than unity, and completing the combustionreaction comprises mixing combustion products from the combustionsection using the vanes of the first set of vanes.
 11. The method ofclaim 8, wherein the engine is operative such that cooling air providedto the first set of vanes reacts with combustion products from thecombustion section in order to complete the combustion reaction.
 12. Themethod of claim 8, further comprising film-cooling the vanes of thefirst set of vanes.
 13. The method of claim 12, wherein: the gas turbineengine further comprises a second set of vanes located downstream fromthe first set of blades; and the second set of vanes is film-cooled to alesser extent than that provided to the first set of vanes.